1. Field of the Invention
The present invention generally relates to rigid solar panels, and more specifically to a lightweight panel that includes a solar cell carrier, which also functions as a heat sink, mounted on a rigid lattice structure that provides both axial and bending stiffness.
2. Description of the Related Art
As shown in FIG. 1, a spacecraft 10 includes a number of rigid solar panels 12, which are shown in their deployed position. Solar arrays 14 which may include hundreds or thousands of solar cells 16 bonded to each solar panel 12 are used to provide electrical power to drive a variety of spacecraft systems and to recharge its batteries. The spacecraft rotates the solar panels 12 so that they receive direct illumination from the sun 18 to increase efficiency.
The solar cells 16 include a flat photovoltaic wafer made from n-type or p-type crystalline semiconductor material, such as silicon, gallium-arsenide or germanium in or on which a thin surface layer of the opposite conductivity type is formed. The interface between the surface layer and the main or bulk region of the wafer defines a semiconductor junction. Illumination of the thin surface layer causes a liberation of charge carriers, including electrons and holes in the region of the semiconductor junction, which migrate toward opposite surfaces to establish a potential across the solar cell.
The solar panel 12 has three primary functions. First, the panel provides a rigid support structure with sufficient axial and bending stiffness for carrying the solar cell array 14 through a dynamically active launch environment into orbit and positioning it to receive illumination. Secondly, the front surface of the solar panel 12 to which the array is bonded is electrically inert so that the individual solar cells 16 are electrically isolated. Lastly, the solar panel 12 serves as a heat sink to the space facing side (opposite sun 18) for the solar cell array.
There are a number of important issues associated with solar panel design. The heat sink capabilities of the panel must be sufficient to cool the solar cell array to maintain power efficiency. Known silicon solar cell arrays operate at approximately 12.5% (beginning of life) efficiency. The solar panel must have high axial and bending stiffness to provide a rigid support structure. The solar panel should have low and matched thermal expansion properties. The temperature on the illuminated side of the array can be as high as +55.degree. C. and can be as low as -160.degree. C. on the back surface, which faces deep space. Absent the desired thermal expansion properties, these temperature gradients could cause significant warping of the solar panel.
A paramount concern in solar panel design is weight. Existing spacecraft have eight solar panels, four per side, where the structural constituents weigh approximately 16.5 lbs (7.48 kg) per panel. Currently, the cost of flying a spacecraft can be estimated as high as $20,000 per pound over the lifetime of the spacecraft. Hence, the weight of the solar panel impacts the overall cost of operating a spacecraft.
FIG. 2 shows a section view of a known implementation of the rigid solar panel 12. The panel is made of a sandwich of two Kevlar.RTM. face sheets 20 and 22 bonded to opposite sides of an aluminum honeycomb core 24 and is 85" by 100" (2.16 m.times.2.54 m) in size. This integrated structure provides both a carrier for the solar cells 16 and a rigid support structure. The Kevlar.RTM. face sheets have a low thermal coefficient of expansion, and when bonded with equal thickness on either side of core 24 create a thermally "matched" structure. Furthermore, Kevlar.RTM. is electrically inert, and thus electrically isolates solar cells 16.
The panel's axial stiffness (S), preferably 1.9.times.10.sup.6 lbs, is given by: EQU S=2T.times.W.times.E (1)
where T is the thickness of each face sheet 20 and 22, W is a normalized width, and E is the material's modulus of elasticity, which for Kevlar.RTM. is 3.5 million pounds per square inch (psi). To achieve adequate axial stiffness, the thickness T of each Kevlar.RTM. face sheet is chosen to be 0.0056 inches (0.14 mm). The two face sheets together weigh approximately 5.5 lbs (2.49 kg).
The panels's bending stiffness (B), preferably 0.96 lbxin.sup.2, is given by: ##EQU1## where D is the distance between face sheets 20 and 22. To provide sufficient bending stiffness, the sheets 20 and 22 must be separated by at least approximately 1 inch (2.54 cm), and hence the honeycomb core 24 has that thickness.
Besides providing the rigid support structure, the honeycomb core 24 also serves as the panel's heat sink. The thermal conductivity (Q) of the core, preferably 8.8 Btu-Hr-.degree.F., is given by: ##EQU2## where k is the core material's coefficient of thermal conductivity in (80Btu hr .degree.F.)/ft, A is the cross-sectional area of the honeycomb in ft.sup.2 and L is the thickness of the core in ft. Aluminum is the preferred core material because it is lightweight and has a high coefficient of thermal conductivity.
FIG. 3 shows the core's honeycomb pattern, which is chosen to both reduce the core's weight and conduct heat away from the solar cells to deep space. The size of the individual honeycomb cells 26 and their wall thickness 28 determine the cross-sectional area A. The core's thermal conductivity Q can be increased by reducing the solar cell size, which effectively increases A. However, this would significantly increase the weight of the core. Hence, the known core uses a 3/8 inch (0.95 cm) cell and has a weight of approximately 4.9 lbs (2.2 kg).